Most relatively small missiles in use today are propelled by solid fuel rockets as opposed to, for example, turbojet engines. The selection of a solid fuel rocket as a propulsion device has been largely dictated by two factors. First, in many instances, a turbine engine cannot be fabricated sufficiently economically as to compete with a solid fuel rocket engine. Secondly, in small size missiles, i.e., those having relatively small diameter on the order of about six inches, it is heretofore been quite difficult to manufacture an efficient turbojet engine that would fit within the six inch envelope required of the propulsion unit for such a missile.
As a consequence of the use of solid fuel rocket engines, some degree of control of the missile flight path or trajectory is lost over that which would be available were it possible to propel the missile by a gas turbine engine whose output can be readily varied. Further, even if the gas turbine engine operates relatively inefficiently, the use of such an engine greatly extends the range of the missile.
The difficulty in economically producing small diameter gas turbine engines resides not so much in the manufacture of the compressor and/or turbine section of the engine, but rather, is more apt to be attributable to the labor intensive nature of the manufacture of the combustor. Furthermore, as combustor sizes shrink to fit within some desired small envelope as the six inch envelope of a relatively small missile mentioned previously, the difficulty in achieving efficient combustion of fuel rises asymptotically. In particular, as the size or volume of a combustor is reduced, there may be insufficient volume to allow the fuel to be first vaporized completely, burned efficiently, and then mixed uniformly.
The present invention is directed to overcoming one or more of the above problems.